Alloy, protective layer and component

ABSTRACT

Known protective layers having a high Cr-content and a silicone in addition, form brittle phases that embrittle further under the influence of carbon during use. The protective layer according to the invention is composed of 22% to 26% cobalt (Co), 10.5% to 12% aluminum (Al), 0.2% to 0.4% Yttrium (Y) and/or at least one equivalent metal from the group comprising Scandium and the rare earth elements, 15% to 16% chrome (Cr), optionally 0.3% to 1.5% tantal, the remainder nickel (Ni).

CROSS-REFERENCE TO RELATED APPLICATIONS

The present application is a 35 U.S.C. §§ 371 national phase conversionof PCT/EP2012/065219, filed Aug. 3, 2012, which claims priority ofEuropean Patent Application No. 11180885.3, filed Sep. 12, 2011, thecontents of which are incorporated by reference herein. The PCTInternational Application was published in the German language.

TECHNICAL FIELD

The invention relates to an alloy, a protective layer for protecting acomponent against corrosion and/or oxidation, in particular at hightemperatures, and a component.

TECHNICAL BACKGROUND

Large numbers of protective layers for metal components, which areintended to increase their corrosion resistance and/or oxidationresistance, are known in the prior art. Most of these protective layersare known by the generic name MCrAlY, where M stands for at least one ofthe elements from the group comprising iron, cobalt and nickel and otheressential constituents are chromium, aluminum and yttrium.

Typical coatings of this type are known from U.S. Pat. Nos. 4,005,989and 4,034,142.

The endeavor to increase the intake temperatures both in static gasturbines and in aircraft engines is of great importance in thespecialist field of gas turbines, since the intake temperatures areimportant determining quantities for the thermodynamic efficienciesachievable with gas turbines. Intake temperatures significantly higherthan 1000° C. are possible when using specially developed alloys as basematerials for components to be heavily loaded thermally, such as guidevanes and rotor blades, in particular by using single-crystalsuperalloys. To date, the prior art permits intake temperatures of 950°C. or more for static gas turbines and 1100° C. or more in gas turbinesof aircraft engines.

Examples of the structure of a turbine blade with a single-crystalsubstrate, which in turn may be complexly constructed, are disclosed byWO 91/01433 A1.

While the physical loading capacity of the base materials so fardeveloped for the components to be heavily loaded is substantiallyunproblematic in respect of possible further increases in the intaketemperatures, it is necessary to resort to protective layers in order toachieve sufficient resistance against oxidation and corrosion. Besidessufficient chemical stability of a protective layer under theaggressions which are to be expected from exhaust gases at temperaturesof the order of 1000° C., a protective layer must also have sufficientlygood mechanical properties, not least in respect of the mechanicalinteraction between the protective layer and the base material. Inparticular, the protective layer must be ductile enough to be able toaccommodate possible deformations of the base material and not crack,since points of attack would thereby be provided for oxidation andcorrosion. The problem then typically arises that increasing theproportions of elements such as aluminum and chromium, which can improvethe resistance of a protective layer against oxidation and corrosion,leads to a deterioration of the ductility of the protective layer sothat mechanical failure is possible, in particular the formation ofcracks, under a mechanical load conventionally occurring in a gasturbine.

SUMMARY OF THE INVENTION

It is therefore an object of the invention to provide an alloy and aprotective layer, having good high-temperature resistance to corrosionand oxidation, and also good longterm stability and which is adaptedparticularly well to a mechanical load which is to be expectedparticularly in a gas turbine at a high temperature.

The object is achieved by an alloy and a protective layer as disclosedherein.

It is another object of the invention to provide a component which hasincreased protection against corrosion and oxidation.

The object is likewise achieved by a component of a gas turbine or steamturbine, which comprises a protective layer of the type described abovefor protection against corrosion and oxidation at high temperatures.

BRIEF DESCRIPTION OF THE FIGURES

The figures and the description merely represent exemplary embodimentsof the invention.

FIG. 1 shows a layer system with a protective layer,

FIG. 2 shows compositions of superalloys,

FIG. 3 shows a gas turbine,

FIG. 4 shows a turbine blade and

FIG. 5 shows a combustion chamber.

DESCRIPTION OF EMBODIMENTS

According to the invention, a protective layer 7 (FIG. 1) for protectinga component against corrosion and oxidation at a high temperatureconsisting essentially of or at least of the following elements(proportions indicated in wt %) and in an alloy:

nickel,

Co: 22%-26%,

Cr: 15%-16%,

Al: 10.5%-12%,

0.2%-0.6% rare earth element (yttrium, . . . ) and/or scandium (Sc):

optionally

Ta: 0.5%-1.5%.

The list of the alloying elements Ni, Co, Cr, Al, Y, Ta is preferablynot conclusive.

Nickel preferably forms the matrix.

The list of Ni, Co, Cr, Al, Y, Ta is preferably conclusive.

The contents of the alloying elements Co, Cr, Al, Y have the followingadvantages:

Moderately High Co Content:

Extension of the beta/gamma field, avoidance of brittle phases such as,for example, the alpha phases, which is deliberately reduced and isnormally regarded as positive for ductility.

Moderate Cr Content:

Sufficiently high for increasing the activity of Al for the Al₂O₃formation;

low enough to avoid brittle phases (alpha chromium or sigma phase).

Moderately High Al Content:

Sufficiently high for Al activity for the formation of a stable Al₂O₃layer;

low enough to avoid embrittlement effects.

Low Y Content:

Sufficiently high to still form sufficient Y aluminate for the formationof Y-containing “pegs” with low oxygen contamination; low enough tonegatively accelerate the oxide layer growth of the Al₂O₃ layer.

Tantalum has a positive effect on the phase stability of the γ′ phase orshifts the transition to higher temperatures and thus slows down thephase degradation by the consumption of aluminum in the layer.

It is to be noted that the proportions of the individual elements arespecially adapted with a view to their effects, which are to be seenparticularly in connection with the element silicon. If the proportionsare dimensioned so that no silicon precipitates are formed, thenadvantageously no brittle phases are created during use of theprotective layer so that the operating time performance is improved andextended.

This is achieved not only by a low chromium content but also, takinginto account the effect of aluminum on the phase formation, byaccurately dimensioning the aluminum content.

In conjunction with the reduction of the brittle phases, which have adetrimental effect particularly with high mechanical properties, thereduction of the mechanical stresses due to the selected nickel contentimproves the mechanical properties.

With good corrosion resistance, the protective layer has particularlygood resistance against oxidation and is also distinguished byparticularly good ductility properties, so that it is particularlyqualified for use in a gas turbine 100 (FIG. 3) with a further increasein the intake temperature. During operation, embrittlement scarcelytakes place since the layer comprises hardly any chromium-siliconprecipitates, which become embrittled in the course of use.

The powders are for example applied by plasma spraying (APS, LPPS, VPS,etc.) in order to form a protective layer. Other methods may likewise beenvisaged (PVD, CVD, SPPS, etc.).

The described protective layer 7 also acts as a layer which improvesadhesion to the superalloy.

Further layers, in particular ceramic thermal barrier layers 10, may beapplied onto this protective layer 7.

In a component 1, the protective layer 7 is advantageously applied ontoa substrate 4 made of a nickel-based or cobalt-based superalloy (FIG.2).

Compositions of this type are known as casting alloys for the substrateunder the references GDT222, IN939, IN6203 and Udimet 500. Otheralternatives for the substrate 4 (FIG. 2) of the component 1, 120, 130,155 are listed in FIG. 2.

The thickness of the protective layer 7 on the component 1 is preferablydimensioned with a value of between about 100 μm and 300 μm.

The protective layer 7 is particularly suitable for protecting thecomponent 1, 120, 130, 155 against corrosion and oxidation while thecomponent is being exposed to an exhaust gas at a material temperatureof about 950° C., or even about 1100° C. in aircraft turbines.

The protective layer 7 according to the invention is thereforeparticularly qualified for protecting a component of a gas turbine 100,in particular a guide vane 120, rotor blade 130 or a heat shield element155, which is exposed to hot gas before or in the turbine of the gasturbine 100 or of the steam turbine.

The protective layer 7 may be used as an overlay (the protective layeris the outermost layer) or as a bondcoat (the protective layer is aninterlayer).

FIG. 1 shows a layer system 1 as a component.

The layer system 1 has a substrate 4.

The substrate 4 may be metallic and/or ceramic. Particularly in the caseof turbine components, for example turbine rotor blades 120 (FIG. 4) orguide vanes 130 (FIGS. 3, 4), heat shield elements 155 (FIG. 5) or otherhousing parts of a steam or gas turbine 100 (FIG. 3), the substrate 4has a nickel-, cobalt- or iron-based superalloy, in particular itconsists thereof.

Nickel-based superalloys (FIG. 2) are preferably used.

The protective layer 7 according to the invention is provided on thesubstrate 4.

This protective layer 7 is preferably applied by plasma spraying (VPS,LPPS, APS, etc.).

It may be used as an outer layer (not shown) or interlayer (FIG. 1).

Preferably, there will be a ceramic thermal barrier layer 10 on theprotective layer 7.

Preferably, the layer system consists of substrate 4, protective layer 7and ceramic thermal barrier layer 10, optionally of a TGO beneath thethermal barrier layer 10.

The protective layer 7 may be applied onto newly produced components andrefurbished components.

Refurbishment means that components 1 are separated if need be fromlayers (thermal barrier layer) after their use and corrosion andoxidation products are removed, for example by an acid treatment (acidstripping). It may sometimes also be necessary to repair cracks. Such acomponent may subsequently be recoated, since the substrate 4 is veryexpensive.

FIG. 3 shows a gas turbine 100 by way of example in a partiallongitudinal section.

The gas turbine 100 internally comprises a rotor 103, which will also bereferred to as the turbine rotor, mounted so as to rotate about arotation axis 102 and having a shaft 101.

Successively along the rotor 103, there are an intake manifold 104, acompressor 105, an e.g. toroidal combustion chamber 110, in particular aring combustion chamber, having a plurality of burners 107 arrangedcoaxially, a turbine 108 and the exhaust manifold 109.

The ring combustion chamber 110 communicates with an e.g. annular hotgas channel 111. There, for example, four successively connected turbinestages 112 form the turbine 108. Each turbine stage 112 is formed forexample by two blade rings. As seen in the flow direction of a workingmedium 113, a guide vane row 115 is followed in the hot gas channel 111by a row 125 formed by rotor blades 120.

The guide vanes 130 are fastened on an inner housing 138 of a stator 143while the rotor blades 120 of a row 125 are fitted on the rotor 103, forexample by means of a turbine disk 133. Coupled to the rotor 103, thereis a generator or a work engine (not shown).

During operation of the gas turbine 100, air 135 is taken in andcompressed by the compressor 105 through the intake manifold 104. Thecompressed air provided at the turbine-side end of the compressor 105 isdelivered to the burners 107 and mixed there with a fuel. The mixture isthen burnt to form the working medium 113 in the combustion chamber 110.From there, the working medium 113 flows along the hot gas channel 111past the guide vanes 130 and the rotor blades 120. At the rotor blades120, the working medium 113 expands by imparting momentum, so that therotor blades 120 drive the rotor 103 and the work engine coupled to it.

The components exposed to the hot working medium 113 experience thermalloads during operation of the gas turbine 100.

Apart from the heat shield elements lining the ring combustion chamber110, the guide vanes 130 and rotor blades 120 of the first turbine stage112, as seen in the flow direction of the working medium 113, are heatedthe most.

In order to withstand the temperatures prevailing there, they may becooled by means of a coolant.

The substrates may likewise comprise a directional structure, i.e. theyare single-crystal (SX structure) or comprise only longitudinallydirected grains (DS structure).

Iron-, nickel- or cobalt-based superalloys are for example used as thematerial for the components, in particular for the turbine blades 120,130 and components of the combustion chamber 110.

Such superalloys are known for example from EP 1 204 776 B1, EP 1 306454, EP 1 319 729 A1, WO 99/67435 or WO 00/44949.

The guide vanes 130 comprise a guide vane root (not shown here) facingthe inner housing 138 of the turbine 108, and a guide vane head lyingopposite the guide vane root. The guide vane head faces the rotor 103and is fixed on a fastening ring 140 of the stator 143.

FIG. 4 shows a perspective view of a rotor blade 120 or guide vane 130of a turbomachine, which extends along a longitudinal axis 121.

The turbomachine may be a gas turbine of an aircraft or of a power plantfor electricity generation, a steam turbine or a compressor.

The blade 120, 130 comprises, successively along the longitudinal axis121, a fastening zone 400, a blade platform 403 adjacent thereto as wellas a blade surface 406 and a blade tip 415.

As a guide vane 130, the vane 130 may have a further platform (notshown) at its vane tip 415.

A blade root 183 which is used to fasten the rotor blades 120, 130 on ashaft or a disk (not shown) is formed in the fastening zone 400.

The blade root 183 is configured, for example, as a hammerhead. Otherconfigurations as a firtree or dovetail root are possible.

The blade 120, 130 comprises a leading edge 409 and a trailing edge 412for a medium which flows past the blade surface 406.

In conventional blades 120, 130, for example solid metallic materials,in particular superalloys, are used in all regions 400, 403, 406 of theblade 120, 130.

Such superalloys are known for example from EP 1 204 776 B1, EP 1 306454, EP 1 319 729 A1, WO 99/67435 or WO 00/44949.

The blade 120, 130 may in this case be manufactured by a casting method,also by means of directional solidification, by a forging method, by amachining method or combinations thereof.

Workpieces with a single-crystal structure or single-crystal structuresare used as components for machines which are exposed to heavymechanical, thermal and/or chemical loads during operation.

Such single-crystal workpieces are manufactured, for example, bydirectional solidification from the melts. These are casting methods inwhich the liquid metal alloy is solidified to form a single-crystalstructure, i.e. to form the single-crystal workpiece, or isdirectionally solidified.

Dendritic crystals are in this case aligned along the heat flux and formeither a rod crystalline grain structure (columnar, i.e. grains whichextend over the entire length of the workpiece and in this case,according to general terminology usage, are referred to as directionallysolidified) or a single-crystal structure, i.e. the entire workpiececonsists of a single crystal. It is necessary to avoid the transition toglobulitic (polycrystalline) solidification in these methods, sincenondirectional growth will necessarily form transverse and longitudinalgrain boundaries which negate the beneficial properties of thedirectionally solidified or single-crystal component.

When directionally solidified structures are referred to in general,this is intended to mean both single crystals which have no grainboundaries or at most small-angle grain boundaries, and also rod crystalstructures which, although they do have grain boundaries extending inthe longitudinal direction, do not have any transverse grain boundaries.These latter crystalline structures are also referred to asdirectionally solidified structures.

Such methods are known from U.S. Pat. No. 6,024,792 and EP 0 892 090 A1.

The blades 120, 130 may also have layers 7 according to the inventionprotecting against corrosion or oxidation.

The density is preferably 95% of the theoretical density.

A protective aluminum oxide layer (TGO=thermally grown oxide layer) isformed on the MCrAlX layer (as an interlayer or as the outermost layer).

On the McrAlX layer, there may furthermore be a thermal barrier layer,which is preferably the outermost layer and consists for example ofZrO₂, Y₂O₃—ZrO₂, i.e. it is not stabilized or is partially or fullystabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide.

The thermal barrier layer covers the entire MCrAlX layer.

Rod-shaped grains are produced in the thermal barrier layer by suitablecoating methods, for example electron beam deposition (EB-PVD).

Other coating methods may be envisaged, for example atmospheric plasmaspraying (APS), LPPS, VPS or CVD. The thermal barrier layer may compriseporous, micro- or macro-cracked grains for better thermal shockresistance. The thermal barrier layer is thus preferably more porousthan the MCrAlX layer.

The blade 120, 130 may be designed to be hollow or solid. If the blade120, 130 is intended to be cooled, it will be hollow and optionally alsocomprise film cooling holes 418 (indicated by dashes).

FIG. 5 shows a combustion chamber 110 of the gas turbine 100. Thecombustion chamber 110 is designed for example as a so-called ringcombustion chamber in which a multiplicity of burners 107, which produceflames 156 and are arranged in the circumferential direction around arotation axis 102, open into a common combustion chamber space 154. Tothis end, the combustion chamber 110 as a whole is designed as anannular structure which is positioned around the rotation axis 102.

In order to achieve a comparatively high efficiency, the combustionchamber 110 is designed for a relatively high temperature of the workingmedium M, of about 1000° C. to 1600° C. In order to permit acomparatively long operating time even under these operating parameterswhich are unfavorable for the materials, the combustion chamber wall 153is provided with an inner lining formed by heat shield elements 155 onits side facing the working medium M.

Owing to the high temperatures inside the combustion chamber 110, acooling system may also be provided for the heat shield elements 155 orfor their retaining elements. The heat shield elements 155 are thenhollow, for example, and optionally also have cooling holes (not shown)opening into the combustion chamber space 154.

Each heat shield element 155 made of an alloy is equipped with aparticularly heat-resistant protective layer (MCrAlX layer and/orceramic coating) on the working medium side, or is made of refractorymaterial (solid ceramic blocks).

These protective layers 7 may be similar to the turbine blades. On theMCrAlX, there may furthermore be an e.g. ceramic thermal barrier layerwhich consists for example of ZrO₂, Y₂O₃—ZrO₂, i.e. it is not stabilizedor is partially or fully stabilized by yttrium oxide and/or calciumoxide and/or magnesium oxide.

Rod-shaped grains are produced in the thermal barrier layer by suitablecoating methods, for example electron beam deposition (EB-PVD).

Other coating methods may be envisaged, for example atmospheric plasmaspraying (APS), LPPS, VPS or CVD. The thermal barrier layer may compriseporous, micro- or macro-cracked grains for better thermal shockresistance.

Refurbishment means that turbine blades 120, 130 or heat shield elements155 may need to be stripped of protective layers (for example bysandblasting) after their use. The corrosion and/or oxidation layers orproducts are then removed. Optionally, cracks in the turbine blade 120,130 or heat shield element 155 are also repaired. The turbine blades120, 130 or heat shield elements 155 are then recoated and the turbineblades 120, 130 or heat shield elements 155 are used again.

The invention claimed is:
 1. A protective single layer for protecting acomponent against corrosion and/or oxidation, the component having asubstrate which is nickel-based or cobalt-based and the single layerresides on the substrate, wherein the single layer has the compositionof an alloy, which consists of: 22% wt % to 26% wt % cobalt (Co), 15% wt% to 16% wt % chromium (Cr), 10.5% wt % to 12% wt % aluminum (Al), atleast 0.3% wt % of tantalum and a maximum of 1.5% wt % tantalum, a totalof 0.2% wt % to 0.6% wt % of yttrium (Y), and the balance nickel (Ni),wherein no silicon precipitates are formed during use at least at 950°C. or at least at 1100° C., thereby preventing embrittlement of thealloy during use at least at 950° C. or at least at 1100° C., and theprotective single layer is applied by plasma spraying, wherein the Cocontent extends the beta/gamma field and avoids the alpha phases atleast at 950° C. or at least at 1100° C., wherein the Cr contentincreases the activity of Al for the Al₂O₃ formation and avoids alphachromium or sigma phase at least at 950° C. or at least at 1100° C.,wherein the Al content forms a stable Al₂O₃ layer and avoidsembrittlement at least at 950° C. or at least at 1100° C., wherein the Ycontent forms sufficient Y aluminate for the formation of Y-containingpegs with low oxygen contamination and negatively accelerates the oxidelayer growth of the Al₂O₃ layer at least at 950° C. or at least at 1100°C., wherein the Ta content stabilizes the γ′ phase or shifts thetransition to a higher temperature to slow down phase degradation byconsumption of aluminum at least at 950° C. or at least at 1100° C., andwherein the single layer protects the substrate against corrosion andoxidation at least at 950° C. or at least at 1100° C. and is ductileenough to be able to accommodate possible deformations of the substrateand not crack to avoid providing points of attack for oxidation andcorrosion at least at 950° C. or at least at 1100° C.
 2. The protectivesingle layer as claimed in claim 1, wherein the alloy consists of 0.5wt. % yttrium (Y).
 3. A component of a gas turbine, having a substrateof the component which is nickel-based or cobalt-based, and theprotective single layer as claimed in claim 1 thereon, configured toprotect against corrosion and oxidation at high temperatures.
 4. Thecomponent as claimed in claim 3, further comprising a ceramic thermalbarrier layer applied onto the protective single layer.
 5. Theprotective single layer as claimed in claim 1, wherein the alloyconsists of at least 0.5 wt. % of tantalum.